Rocket propelled missile



Jan. 26, 1965 'w. CZERWINSKI ETAL 3,167,016-

ROCKET PROPELLED MISSILE Filed July so, 1956 7 Sheets-Sheet 1 g 0 mm m(D m w (5 m m 00 N I N c- (0 E N N v g. N

INVENTORS W CZERWINSKI JVACHAMBERLIN FAWOODWARD BY g ATTORNEYS w.CZERWlNSKl ETAL 3,167,016

Jan. 26, 1965 ROCKET PROPELLED MISSILE 7 Sheets-Sheet 2 Filed July 50,1956' NO vm mm mm on mm mm hm @N Wm OmnN N05 mm 5mm Om hm MD. a m w m wm .NIWMF. w

ATTORNEYS Jan. 26, 1965 w. CZERWINSKI ETAL 3,167,016

ROCKET PROPELLED MISSILE Filed July 30, 1956 7 Sheets-Sheet 3 INVENTORSW CZERWINSKI JACHAMBERLIN F .A.WOODWARD ATTORNEYS Jan. 26, 1965 w.CZERWINSKI ETAL 3,167,016

ROCKET PROPELLED MISSILE Filed July 30, 1956 7 Sheets-Sheet 4 INVENTORSWCZERWINSKI J.A.CHAMBERLIN FAWOODWARD BY%; z &

ATTORNEYS Jan. 26, 1965 w. CZERWINSKI ETAL 3,167,015

ROCKET PROPELLED MISSILE Filed July so, 1956 7 Sheets-Sheet 5 INVENTORSW CZERWINSKI JACHAMBERLIN F .AWOODWARD ATTORNEYS 1965 w.'czE'Rw|NsK'|ETAL 3,167,016

ROCKET PROYPQELLED MISSILE 7 Sheets$heet 6 Filed July 30, 1956 Nm vmpmmp 0MP mmw INVENTORS WCZERWINSKI J.A.CHAMBERLIN F.A.WOODWARD ATTORNEYSJan. 26, 1965 w. CZERWINSKI ETAL ,0

ROCKET PROPELLED MISSILE 7 Sheets-Sheet 7 Filed July 30, 1956 m: Om GNPm mm 09 D: MNF vmF DNF .I1.l|.||

United States Patent 3,167,016 RGCKET IRGPELLED MISSILE WaciawCzerwinski, James Arthur Chamherlin, and

Frank Arthur Woodward, all of Toronto, Ontario, Canaria, assignors, bymesne assignments, to The De Havil- Band Aircraft of Canada Limited,Downsview, ()ntario,

Canada, a corporation Filed July 30, 1956, Ser. No. 600,859 8 Claims.(Cl. 102-49) This invention relates to missiles and more particularly toan unguided two stage missile propelled by the efilux of gases generatedby a rocket motor.

In the past, surface installations either on land or on the sea, havebeen defended against air launched attacks by the use of variousspecialized devices, the most common being projectiles propelled throughthe rifled barrel of a gun directed at the target. Other devices includeunguided rocket projectiles, fired either singly or in salvo, andelectronically guided missiles which are directed to the vicinity of thetarget by external aerodynamic control surfaces.

Unguided projectiles have the characteristic of low single shotprobability of kill, which has been partially overcome by firing highvelocity projectiles fitted with proximity fuses, at a high rate fromradar directed guns or projectors. However, in view of the advent offast modern aircraft which can take extreme evasive action, and ofelectronically controlled glide bombs and improved aerial torpedoes, itbecomes necessary to increase substantially the warhead effectiveness ofunguided missiles and to reduce their time of flight if an acceptablekill probability is to be maintained.

Homing guided missile systems have been used with a considerable degreeof success. The guided missile overcomes the disadvantage of a long timeof flight by continuously correcting its path in relation to the targetso that the latters evasive action becomes substantially meaningless.However, such a weapon with its large amount of electronic guidanceequipment requires a long period of time to develop and is veryexpensive to manufacture.

The subject invention embodies an unguided two stage rocket propelledanti-aircraft missile which will achieve an exceptionally high singleshot probability of kill. The missile described herein has a warheadweighing in the order of 150 lbs. and travels with a mean velocity of6,000 ft./sec. over an effective range of up to 6,000 yds. This warheadis about eight times as large as a conventional anti-aircraft shell, andits velocity is about three times as high as that of a conventionalanti-aircraft shell. These features make it possible to achieve anacceptable probability of kill against any of the targets previouslymentioned, against which conventional defensive armament becomesineffective. In addition, the rocket of the invention has the specialfeatures of spin stabilization, which eliminates the necessity ofexternal fins on the outer casing, and thus simplifies ground handling;furthermore it has a simple mechanical construction combined with aminimum of delicate electronic or mechanical systems, thus reducing thecost as compared to that Of a conventional guided missile.

It is obvious that such a missile will have to be aimed in much the sameway as a gun and will require afire control system of similar accuracy.A launching device therefore is provided: it is directed by an automaticelectronic fire control system which will not be described herein sinceit is not part of the invention.

It is an object of the invention to provide a two stage rocket propelledmissile having spin stabilization and then, on separation, finstabilization.

It is another object of the invention to provide a simple 3,167,016Patented Jan. 26, 1965 but elfective mechanical arrangement whereby thetwo stages are separated when the propellant charge in the first stageis exhausted.

Another object of the invention is to provide structure associated withboth the launcher and the missile which ensures that all parts of themissile leave the guiding surfaces of the launcher at one time in orderto minimize misguide due to movement of the launcher during firing.

The foregoing and other objects and advantages of the invention willbecome apparent from a study of the following specification, taken inconjunction with the accompanying drawings, in which like referencenumerals indicate corresponding parts throughout the several views, andin which:

FIGURE 1 is a longitudinal cross-sectional view of a missile constructedin accordance with the invention;

FIGURE 1A is an enlarged longitudinal cross-sectional view of the foreand aft portions of the missile of FIG- URE 1;

FIGURE 2 is a longitudinal, part-elevational and partsectional view ofthe missile at the initiation of separation of the first and secondstage rockets;

FIGURE 2A is an enlarged longitudinal, part-elevational andpart-sectional view of the fore and aft portions of the missile at theinitiation of the separation of the first and second stage rockets;

FIGURE 3 is a transverse sectional view of the missile taken on the line33 of FIGURE 1;

FIGURE 4 is a perspective view of the first and second stage rockets ofthe missile immediately after separation;

FIGURE 5 is a perspective view of a head segment with its associatedfairing segment;

FIGURE 6 is a perspective view of a launcher for the missile, one of thebarrels of the launcher being broken away to show a missile in firingposition in the barrel;.

FIGURE 7 is a longitudinal sectional view through, a barrel of thelauncher showing a missile in firing position;

FIGURE 7A is a detail of FIGURE 7 on a larger scale and showing a stopfor the missile;

FIGURE 8 is an end elevational view of the launcher of FIGURE 6; and

FIGURE 9 is a view similar to FIGURE 7 but showing the missile emergingfrom the barrel.

Referring particularly to FIGURES 1, 2, 1A and 2A a missile 20 comprisesa first stage 21, and a second stage 22 coaxially disposed within andsubstantially encased by the first stage. The first stage includes anouter shell or casing 23 of cylindrical cross-section, to the aftportion of which is welded or otherwise secured a housing 24 of similarcylindrical cross-section. The housing is, in the vicinity of itsattachment to the shell 23, of slightly larger diameter than the shell23 and has machined a recess 25 in its inner cylindrical surfaceadjacent its forward edge. The aft portion of the housing is of greaterdiameter than the forward portion, to provide a rear supporting ring 26the main function of which will become apparent later. The ring 26 isprovided with tapped holes 27 into which bolts 28 are inserted to attacha nozzle assembly 29 comprising an outer ring 30 and an inner ring 31,the space therebetween constituting an annular convergentdivergentnozzle 32 in which are disposed deflector vanes 33. The vanes 33 deflectthe exhaust gases through an angle of approximately 20 from the axis toimpart an angular velocity or spin to the rocket. The vanes are soarranged that the thrust is directed toward the centre of gravity of themissile in order to minimize the effects of malalignment due to thrustdeviations around the nozzle. The vanes 33 are fast to the rings 30 and31 and the vanes and the rings together with a spherical segment 34which encloses the space within the inner ring 31 constitute a rigidunit.

The propellant charge for the first stage rocket comprises amultiplicity of precast rings or grains which are manufactured from somesuitable rapid burning rocket fuel. It will be noted that the innerdiameters of the rings of propellant 35 gradually increase from fore toaft where, at the final ring, the grain assembly is held in position bya segmented rear grain support 36 having flanges 37 whereby it isrotatably mounted in the recess 25. A forward grain support 38 servesmainly to retain the propellant charge in position during loading andstorage.

The inner wall of the first stage rocket is formed by a slendercylindrical shell 39 which also provides the outer casing of the secondstage rocket 22. The second stage rocket is coaxially disposed withinthe first stage rocket and terminates at its aft and in aconvergent-divergent nozzle 40 of circular cross-section which is weldedto the shell 39. The discharge portion of the nozzle 40 mates at itsouter edge with the inner ring 31 of the nozzle 32 and, together with asplit retaining ring 41, forms a gastight joint between the nozzles 32and 40.

A fin wrapper 42 is constituted by a body portion 43 to which areattached integral fins 44 arranged in a slight spiral. The fin wrapperis fastened to the nozzle 40 in torque transmitting relationship bymeans of fore and aft splines 45 and 46'machined externally on the outersurface of the nozzle 40; the splines are complementary to and registerwith splines 47 and 48 respectively, which are machined internally atthe ends of the body portion 43, Thesplines 46 also engage splines 49machined on the inner riiig 31 of the first stage nozzle assembly 29 toprovide a torque transmitting connection between the two stages. Agroove 50 machined in the forward face of the body portion 43 mates witha groove 51 machined in the nozzle 40, the two mating grooves beingadapted to receive a corrugated wire spring 52. Grooves 53 and 54machinedin the splines 46 and 48 respectively, receive the splitretaining ring 41. Notches 55 provided between the segments of the firststage rear grain support 36 allow the passage of the fins 44 onseparation of the two stages.

Near the forward end of the shell 39 (see FIGURES 1A and 2A) is a. breakin which a retaining ring 56 is welded in place. Somewhat forwardly ofthe ring 56 a stepped ring 57 is welded to the exterior of the shell 39for a purpose that will become apparent hereafter. The extreme forwardend of the shell 39 terminates in a forward tube collar 58 which isprovided with an integral collet 59 in which is held a forward grainsupport 60. The grain support 60 positions the second stage propellantcharge, which charge is constituted by rings or grains 61. To the outerperiphery of a dished annular sealing plate 62 is welded an outer collar63 having a groove 64 which receives a seal 65. A groove 66 machined inthe forward tube collar 58 receives a split retaining ring 67 whichholds in position the outer collar 63 and hence the sealing plate 62. Aninner collar 69 extends through the hole 68 inthe centre of the sealingplate and it is welded to the sealing plate. From the aft side of thecollar 69 projects an igniter 70 for the second stage grain; a warhead71 is attached to the other side of the collar 69 by means of screwthreads.

Referring again to the first stage (see FIGURES 1 and 1A) there iswelded to the forward end of the shell 23 a forward ring 72 having,integral with its outer surface, eight winglets 73. Spanning the annularspace between the forward ring 72 and the retaining ring 56 are eighthead segments generally indicated at 74 (see FIGURES 1A and 5); they areannularly arranged about the retaining ring. Each of the head segments74 comprises a web 75 having inner and outer arcuate flange segments 76and 77 respectively, and radially disposed edges 78 which abut the edgesof adjacent webs to form gas tight joints. Sealing grooves 79 and 80 aremachined in the inner and outer flange segments respectively to receiveO-ring seals 81 and 82 respectively. To the forward faces of each of thewebs 75 are attached two integral ribs 83 having toe portions 84 whichare restrained by the stepped ring 57. A notch 85 on each of the toeportions provides an area of high local stress concentration whensuitable loads are applied to the ribs 83. A split retaining ring 86fits into a groove 87 machined on the forward part of the inner surfaceof the forward ring 72 and mates witha recess 88 in the forward edges ofthe annulus constituted by the arcuate flange segments 77; an igniterring 89 abuts the end of the ring 72. The igniter ring 89, which hashigh electrical conductivity, is held securely in a mounting ring 90 ofU-cross-section and which is made of some heat resistant plasticmaterial having good electrical insulating characteristics. The igniterring 89 is attached to the end of the ring 72 by means of an integralflanged collar 91 which fits within the forward inner surface of thering 72 and abuts the split retaining ring 86. Electrical contactbetween the ring 89 and igniters 92 adjacent the first ring ofpropellant is made by means of eight leads 93, each one of which isconnected to a terminal post 94 on the respective eight head segments 74and makes contact with the igniters. The igniters 70 and 92 will not bedescribed in detail since they are of known construction.

Low shock, high heat release, closureless igniters preferably are usedfor reliability. This type of igniter sprays molten magnesium andaluminum powder intimately into the forward disk of propellant; it isreadily available in fully developed form and is known to perform well.

Fairing segments 95, (see FIGURES 2, 2A and 5) when assembled,constitute a fairing which extends from the forward ring 72 to theforward tube collar 58. Each of the fairing segments is provided with apair of stiffening ribs 96, the inner edges of which abut the outeredges of the opposed ribs 83 of the webs 75; in the ribs 96 are holes 97through which pass bolts 98 to fasten the fairing segments to the ribs83.

A warhead assembly 99 (see FIGURE 1A) comprises the warhead 71 having afragmentable casing which contains a high explosive. The warhead 71 isfastened by screw means to the collar 69 and is contained within awarhead fairing 101 shaped in the form of a truncated cone. Secured tothe aft edge of the fairing is a sleeve 1-32; the sleeve includes a face104 which abuts the ends of the fairing segments 95, and a flange 103which is encompassed by the split retaining ring 67. The forward end ofthe fairing 101 is closed by a plate 105 fast to the fairing and havinga threaded flange 106. A hole 107 in the plate 105 allows the plug end108 of a fuse 109 to be screwed into the warhead 71. By the assembly ofthe collar 69, the warhead 71, the fairing 101 and the fuse 109, thisportion of the warhead assembly is secured to the second stage rocket22. A conical tip fairing 110 has at its base a threaded sleeve 111 forattachment to the flange 106. A miniature radar antenna 112 (see FIGURES1 and 2) receives signals from the ground control station which aretransmitted to the fuse 109 to explode the warhead at the mostadvantageous position with respect to the target. The fuse 109, thewarhead 71 and the igniters 70 and 92 will not be described in detailsince they are well-known to those skilled in the art.

In order to ensure that the missile 20 is properly launched, there isprovided a launcher unit 113 (see FIG- URES- 6, 7, 7A, 8 and 9)comprising three separate launching barrels 114 capable of firing threemissiles simultaneously or in a prescribed sequence. Each barrelconsists of a hollow steel cylinder 115 to the exterior of which areintegral stiffening rings 116 which provide circular stability, and alsoaxial flutes 117. In the interior of each barrel (see FIGURES 7 and 9)eight integral missile guide rails 118 are positioned equiangularlyaround the inner surface. In order to achieve dimensional accuracy, itis proposed that the barrels should be manufactured as centrifugal steelcastings. This type of construction not only results in a product havingacceptable dimensional tolerances but also permits the inclusion of allthe exterior reinforcements and the interior guide rails 118.

The three barrels 114 are joined together to form the launcher unit 113by welding the abutting flutes 117 of adjoining barrels; this eliminatesdistortion during welding. The unit 113 is supported on trunnions 119 ofa torque box structure 120 consisting of steel plate of appropriatethickness welded to two suitably positioned stiffening rings. Gearedelevating arcs 121 are welded to the torque box 120 below each of thetrunnions 119. End caps 122 are hingedly mounted on brackets 123; eachend cap may be swung from open to closed position by a rod and crankmechanism 124 welded to a common hinge pin 125 and actuated by ahydraulic jack 126. The end caps may be closed to protect empty orloaded barrels against the weather or, when a rocket has misfired, thecaps may be closed and the missile ejected by compressed air introducedinto the barrel between the missile and the caps. A cutout in the wallof each cylinder 115 (see FIGURES 7 and 7A) receives a missile stop 128hingedly mounted on a pin 129 and urged inwardly by a leaf spring 130. Aspring loaded contact pin 131 positioned immediately forward of one ofthe missile guide rails 118 is connected to the automatic electronicfire control and delivers an electrical impulse to the igniter ring 89when it is desired to launch a missile. The trunnions 119 are mounted ona carriage (not shown) which contains suitable azimuth and elevationgearing to position the launcher unit 113 for firing.

Assembly Because of the specific configuration of the two stages of themissile, one arranged within the other, and because of the novel meansnecessary to effect complete separation, at special sequence of assemblyoperations is required. The first step includes the filling of thesecond stage shell 39 with its propellant 61, held in position by thegrain support 60; then the dished sealing plate 62 together with itsouter collar 63 and its inner collar 69 and the igniter 70 are partiallyinserted into the forward end of the second stage. The seal 65 then isfitted into the groove 64 provided in the collar 63 and the wholesubassembly is completely inserted leaving the groove 66 unobstructed.The split retaining ring 67 finally is placed in the groove 66 tocomplete this subassembly.

The second step in the assembly of the missile is begun by attaching therear grain support 36 to the casing 23 of the first stage 21 and fillingthe casing with the propellant 35. The inner sealing ring 81 then ispositioned around the retaining ring 56 and the eight head segmentsstage sub-assembly is inserted into the first stage sub-assembly and itis pushed as far aft within the first stage sub-assembly as permissiblein order to facilitate the installation of the split retaining ring 8-6in the groove 87. After the ring 86 has been fitted, the second stagesub-assembly is brought forward until the recess 88 in the web 75 of thehead segments 74 abuts the split retaining ring 86.

The third step in the assembly of the missile relates to the nozzlesection (see FIGURES 1A and 2A). The colrugated wire spring 52 islocated in the machined groove 51, and the splines 47 and 48 on the bodyportion 43 of the fin wrapper 42 are pushed over the nozzle splines and46, insuring that the fins 44 slide into the notches of the rear grainsupport 36. The fin wrapper 42 is pushed far enough forward to depressthe corrugated spring 52 and expose the groove 53 in the aft portion ofthe nozzle 40 whereupon the split retaining ring 41 is in- 6 stalled andthe fin wrapper 42 is freed in order to let the face of the groove 54abut the ring 41. The first stage nozzle system 29 is then positioned atthe end of the housing 24 with the inner ring 31 mating with the nozzle40; the nozzle 29 then is bolted to the housing 24.

The fourth and final step involves the assembly of the extreme forwardportion of the missile. The warhead 71 is screwed onto the collar 69after which the warhead fairing 101 is fitted into place with the flange103 of the sleeve 192 being a push fit over the split retaining ring 67;by screwing the fuse 109 into the warhead 71, the fairing 101 is heldfirmly in position. The tip fairing 110 is assembled by screwing thethreaded portion of the sleeve 111 onto the flange 106. Then to completethe assembly, the igniter ring 89 is fitted into position, the leads 93fastened on the face of the webs 75 and the fairing segments 95 placedin position with the bolts 98 fastening them to the ribs 83.

Operation In operation, the end caps 122 of each launching barrel 114are opened and remain open during firing. The forward end of the missile20 is inserted in the barrel and the missile is pushed forward until oneof the winglets 73 engages the missile stop 128, at which stage the pin131 makes contact with the igniter ring 89 (see FIGURES 6, 7 and 7A).The fore part of the missile is supported on the inner surface of thecylinder 115 by the winglets 73 while the aft part is supported by therear ring 26 of the housing 24 on the eight missile guide rails 118. Itis intended that the launcher be aimed by radar control in which casethe command for firing will be automatic, delivering an electricalimpulse to the pin 131. The current will flow through the igniter ring89 and the leads 93 to each of the contacts on the webs 75 of the headsegments 74 to initiate combustion of the first of the grains ofpropellant 35. The hot gases thus formed will, in passing down thediverging passage between the propellant 35. and the shell 39, ignitgthe accelerating compound between the discs of propellant. An inhibitoris provided on the cylindrical surfaces of the discs so that the burningtakes place on the faces to provide a large constant area burningsurface. The gases generated travel radially inwards and then turnthrough 90 to pass axially down the diverging passage. The passage isrearwardly divergent in order to accommodate the increasing volume ofgas as it progresses rearwardly. In order to prevent breaking up of thepropellant under the high axial accelerations encountered, the discs arebonded to the shell 23 and are also supported by the rear grain support36.

The gases, as they pass rearwardly enter the nozzle system 29 and areejected to atmosphere through the annular nozzle 32 to produce a thruston the missile. At the same time, the gases react on the vanes 33 toimpart a spin to the missile, which spin acts as the stabilizing forcewhile the missile is in flight. The rotational forces on the first stagerocket are transferred to the second stage by means of the splinedconnection between the inner ring 31 of the nozzle 32 and the aftportion of the nozzle 40. Because of the high pressure generated by thegas as it is formed in the first stage, high collapsing stresses will beimposed on the shell 39 of the second stage. Therefore, the shell mustbe of sufficient thickness to withstand the high load; it should not beoverlooked however that the second stage grain will aid in withstandingthis radial compressive stress. If it is considered that an increasedshell thickness would seriously affect design criteria, the second stagemay be pressurized immediately prior to use.

As the missile progresses down the launcher, it is supported, as statedabove, by the winglets 73 hearing on the inner surface of the cylinder115 and by the rear ring 26 hearing on the guide rails 118. However, therelationship between the distance from the winglets to the rear ring andthe total length of the launcher barrel relative to the length of theguide rails is such that at the instant 7 the winglets emerge from thelauncher, the rear ring ceases to bear on the guide rails (see FIGURE9). This has the effect of completely freeing the entire missile fromthe launcher at one instant, thus eliminating ballistic dispersion dueto tip ofis, transverse acceleration and static unbalance.

The second stage igniter 70 is activated at the same instant as thefirst stage is fired. However, by means of a time delay fuse, the secondstage propellant is not ignited until a predetermined period of timeafter the firing of the first stage. This period of time is calculatedto coincide with the burn-out time of the first stage propellant. Whenthe pressure in the second stage reaches a predetermined value (calledthe separation pressure), the axial force acting on the sphericalsegment 34 which encloses the second stage nozzle 40, pushes the wholeouter shell 23 aft with sufficient force to break the toe portions 84 ofthe ribs 83 at the notches 85 (see FIGURE The relative movement of theouter shell and the inner shell will cause a swinging movement or radialdispersion of the head segments 74 and of the fairing segments 95attached to them (see FIGURES 2 and 2A). The dispersal speed will beaugmented by the centrifugal force caused by the rotational velocity ofthe missile, as well as by the aerodynamic lift acting on the fairingsegments. During separation, the eight fins 44 at the aft end of thesecond stage will pass freely through the notches 55 in the rear grainsupport. If any difference exists between the relative velocities of thetwo stages at the instant of separation the rear grain support in freeto rotate in the recess 25 in order to obviate any ill efiects due tothe difference in relative velocities. Moreover, during separation, thehead segments 74 with their associated fairing segments 95 will breakaway as shown in FIGURE 4.

At the instant of separation, the first stage becomes, in effect, asecond launcher. Because of the high angular velocity imparted to thefirst stage by reason of the gases discharging through the inclinedvanes 33, the first stage has a high order of gyro stability and ismaintained in a straight path during separation. The second stage isheld concentric with the outer shell 23 by the fins 44 at the rear andalso by the high gas pressure between the shells 23 and 39. The fins 44of the second stage are slightly helical so that the second stageretains the angular velocity imparted to it by the first stage andconsequently retains a certain amount of gyro stability. However,because of the high length to diameter ratio of the second stage anadditional amount of fin stabilization is required and the fins areprovided mainly for this purpose.

After separation, the second stage carrying the warhead 71 continues onits path toward the target. It is not required, nor is it evendesirable, that the warhead be exploded by impact with the target andfor this reason some form of fusing is provided which will automaticallyexplode the warhead at its optimum position in relation to the target.In the embodiment of the invention described herein, it is proposed touse what is known as a Command Fuse. In this type of fuse, a radar rangefinder at the launching base constantly measures the distance betweenthe target and the missile and, at the optimum distance, sends out asignal which is received by the antenna 112 and which triggers the fuse.However, it will be understood that other types of fusing may proveadvantageous for certain applications, e.g. an infra-red fuse operatedby heat generated by the target, a Doppler radar fuse, anaccelerometer-type distance measuring fuse, or an electronicself-contained distance measuring fuse.

It is to be understood that the form of the invention herein shown anddescribed is to be taken as a preferred example of the same and thatvarious changes in the shape, size and arrangement of the parts may beresorted to without departing from the spirit of the invention and thescope of the subjoined claims.

What we claim as our invention is:

1. A rocket propelled missile comprising a first stage rocket motorincluding a cylindrical shell, a second stage rocket motor including acylindrical shell coaxially disposed within and substantially encased bythe first stage motor, the cylindrical shell of the second stage motorbeing spaced from the cylindrical shell of the first stage motor todefine therewith an annular chamber, a solid propellant in the annularchamber and bonded to the inner face of the cylindrical shell of thefirst stage motor, igniting means for the propellant, the solidpropellant being of annular cross-section and its internal diameterbting greater than the external diameter of the shell of the secondstage motor so that an annular passage for products of combustion isprovided between'the propellant and the second stage motor, and'anannular nozzle at the aft end of the shells to discharge to atmospherethe products of combustion emanating from the passage.

2. A rocket propelled missile comprising a first stage rocket motorincluding a cylindrical shell, a second stage rocket motor including acylindrical shell coaxially dis posed within and substantially encasedby the first stage motor, the cylindrical shell of the second stagemotor being spaced from the cylindrical shell of the first stage motorto define therewith an annular chamber, a solid propellant in theannular chamber and bonded to the inner face of the cylindrical shell ofthe first stage motor, igniting means for the propellant, the solidpropellant being of annular cross-section and its internal diameterbeing greater than the external diameter of the shell of the secondstage motor so that an annular passage for products of combustion isprovided between the propellant and the second stage motor, a solidpropellant bonded to the inner face of the cylindrical shell of thesecond'stage motor, the said propellant being of annular cross-sectionto provide a central passage for .the products of combustion thereof,igniting means for the propellant of the second stage motor, and nozzlesat the aft end of the passages to discharge to atmosphere the productsof combustion emanating therefrom.-

3. A rocket propelled missile comprising a first stage rocket motorincluding a cylindrical shell, a second stage rocket motor including acylindrical shell coaxially disposed within and substantially encased bythe first stage motor, the cylindrical shell of the second stage motorbeing spaced from the cylindrical shell of the first stage motor todefine therewith an annular chamber, a solid propellant in the annularchamber and bonded to the inner face of the cylindrical shell of thefirst stage motor, igniting means for the propellant, the solidpropellant being of annular cross-section and its internaldiameter beinggreater than the external diameter of the shell of the second stagemotor so that an annular passage for products of combustion is providedbetween the propellant and the second stage motor, an annular nozzle atthe aft end of the annular passage, a solid propellant bonded to theinner face of the cylindrical shell of the second stage motor, the saidpropellant being of annular cross-section to provide a central passagefor the products of combustion thereof, igniting means for thepropellant of the second stage motor, and a nozzle at the aft end of thepassage of the second stage motor to discharge to atmosphere theproducts of combustion emanating therefrom.

4. A rocket propelled missile comprising a first stage rocket motorincluding a cylindrical shell, a second stage rocket motor including acylindrical shell coaxially disposed within and substantially encased bythe first stage motor, the cylindrical shell of the second stage motorbeing spaced from the cylindrical shell of the first stage motor todefine therewith an annular chamber, a solid propellant in the annularchamber and bonded to the inner face of the cylindrical shell of thefirst stage motor, igniting means for the propellant, the solidpropellant being of annular cross-section and its internal diameterbeing greater than the external diameter of the shell of the secondstage motor so that an annular passage for products of combustion isprovided between the propellant and the second stage motor, and nozzlemeans at the aft end of the first stage motor shell to discharge toatmosphere the products of combustion emanating from the passage.

5. A rocket propelled missile comprising a first stage rocket motorincluding a cylindrical shell, a second stage rocket motor including acylindrical shell coaxially dis posed within and substantially encasedby the first stage motor, the cylindrical shell of the second stagemotor being spaced from the cylindrical shell of the first stage motorto define therewith an annular chamber, a solid propellant in theannular chamber and bonded to the inner face of the cylindrical shell ofthe first stage motor, igniting means for the propellant, the solidpropellant being of annular cross-section and its internal diameterbeing greater than the external diameter of the shell of the secondstage motor so that an annular passage for products of combustion isprovided between the propellant and the second stage motor, nozzle meansat the aft end of the first stage motor and communicating with thepassage for the ejection of the products of combustion of the propellantof the first stage motor to propel the missile and to impart to themissile a rotary motion and thus to provide it with gyroscopicstability, and fin means at the aft end of the shell of the second stagemotor to provide aerodynamic stability thereto after separation from thefirst stage motor.

6. A rocket propelled missile comprising a first stage rocket motorincluding a cylindrical shell, a second stage rocket motor including acylindrical shell coaxially disposed within and substantially encased bythe first stage motor, the cylindrical shell of the second stage motorbeing spaced from the cylindrical shell of the first stage motor todefine therewith an annular chamber, a solid propellant in the annularchamber and bonded to the inner face of the cylindrical shell of thefirst stage motor, igniting means for the propellant, the solidpropellant being of annular cross-section and its internal diameterbeing greater than the external diameter of the shell of the secondstage motor so that an annular passage for the products of combustion isprovided between the propellant and the second stage motor, an annularnozzle at the aft end of the first stage motor and communicating withthe passage for the ejection of the products of combustion of saidpropellant, a series of vanes angularly disposed within the annularnozzle to impart to the missile a rotary motion to provide it withgyroscopic stability before separation of the motors, a nozzle at theaft end of the shell of the second stage motor, and fin means associatedwith said last named nozzle to provide aerodynamic stability to thesecond stage motor after separation thereof from the first stage motor.

7. A rocket propelled missile comprising a first stage rocket motorincluding a cylindrical shell, a second stage rocket motor including acylindrical shell coaxially disposed within and substantially encased bythe first stage motor, the cylindrical shell of the second stage motorbeing spaced from the cylindrical shell of the first stage motor todefine therewith an annular chamber, a solid propellant in the annularchamber and bonded to the inner face of the cylindrical shell of thefirst stage motor, igniting means for the propellant, the solidpropellant being of annular cross-section and its internal diameterbeing greater than the external diameter of the shell of the secondstage motor so that an annular passage for products of combustion isprovided between the propellant and the second stage motor, an annularnozzle at the aft end of the annular passage for discharging theproducts of combustion of the first stage motor to atmosphere, a seriesof vanes angularly disposed within the nozzle to impart to the missile arotory motion and thus to provide it with gyroscopic stability beforeseparation of the motors, a nozzle at the aft end of the second stagemotor for discharging its products of combustion to atmosphere, and agroup of fins equiangularly arranged on the nozzle of the second stagemotor to provide the second stage motor with aerodynamic stability afterseparation of the motors.

8. A missile as claimed in claim 7, in which the fins are helical inshape with their helix axis coinciding with the longitudinal axis of thesecond stage motor.

References Cited by the Examiner UNITED STATES PATENTS 562,535 6/96Hurst l025l X 1,102,653 7/14 Goddard 102-49 X 2,091,635 8/ 37 Hayden102-69 2,246,429 6/ 41 Brandt 102-69 2,478,774 8/49 Meinel 89-1.72,611,317 9/52 Africano 10250 2,686,473 8/54 Vogel 102-49 2,701,984 2/55Terce 10249 2,818,779 1/58 Koeper 89--1.7

SAMUEL FEINBERG, Primary Examiner. SAMUEL BOYD, Examiner.

1. A ROCKET PROPELLED MISSILE COMPRSING A FIRST STAGE ROCKET MOTORINCLUDING A CYLINDRICAL SHELL, A SECOND STAGE ROCKET MOTOR INCLUDING ACYLINDRICAL SHELL COAXIALLY DISPOSED WITHIN AND SUBSTANTIALLY ENCASED BYTHE FIRST STAGE MOTOR, THE CYLINDRICAL SHELL OF THE SECOND STAGE MOTORBEING SPACED FROM THE CYLINDRICAL SHELL OF THE FIRST STAGE MOTOR DEFINETHERWITH AN ANNULAR CHAMBER, A SOLID PROPELLANT IN THE ANNULAR CHAMBERAND BONDED TO THE INNER FACE OF THE CYLINDRICAL SHELL OF THE FIRST STAGEMOTOR, IGNITING MEANS FOR THE PROPELLANT, THE SOLID PROPELLANT BEING OFANNULAR CROSS-SECTION AND ITS INTERNAL DIAMETER BEING GREATER THAN THEEXTERNAL DIAMTERR OF THE SHELL OF THE SECONE STAGE MOTOR SO THAT ANANNULAR PASSAGE FOR PRODUCTS OF COMBUSTION IS PROVIDED BETWEEN THEPROPELLANT AND THE SECOND STAGE MOTOR, AND AN ANNULAR NOZZLE AT THE AFTEND OF THE SHELLS TO DISCHARGE TO ATMOSPHERE THE PRODUCTS OF COMBUSTIONEMANATING FROM THE PASSAGE.